1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to cooling of a turbine airfoil exposed to a high firing temperature.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a hot gas flow is passed through a turbine to extract mechanical energy used to drive the compressor or a bypass fan. The turbine typically includes a number of stages to gradually reduce the temperature and the pressure of the flow passing through. One way of increasing the efficiency of the engine is to increase the temperature of the gas flow entering the turbine. However, the highest temperature allowable is dependent upon the material characteristics and the cooling capabilities of the airfoils, especially the first stage stator vanes and rotor blades. Providing for higher temperature resistant materials or improved airfoil cooling will allow for higher turbine inlet temperatures.
Another way of increasing the engine efficiency is to make better use of the cooling air used that is used to cool the airfoils. A typical air cooled airfoil, such as a stator vane or a rotor blade, uses compressed air that is bled off from the compressor. Since this bleed off air is not used for power production, airfoil designers try to minimize the amount of bleed off air used for the airfoil cooling while maximizing the amount of cooling produced by the bleed off air.
In the industrial gas turbine engine (IGT), high turbine inlet temperatures are envisioned while using low cooling flows. The low cooling flows pass the compressed cooling air through the airfoils without discharging film cooling air out through the airfoil surface and into the hot gas flow or discharging a very minimal amount out through the blade tip. Thus, there is a need for an improvement in the design of low flow cooling circuits for airfoils exposed to higher gas flow temperatures.